Efficient, low pressure ratio propulsor for gas turbine engines

ABSTRACT

A gas turbine engine includes a gear assembly, a bypass flow passage, and a core flow passage. The bypass flow passage includes an inlet. A fan is arranged at the inlet of the bypass flow passage. A first shaft and a second shaft are mounted for rotation about an engine central longitudinal axis. A first turbine is coupled with the first shaft such that rotation of the first turbine is configured to drive the fan, through the first shaft and gear assembly, at a lower speed than the first shaft. The fan includes a hub and a row of fan blades that extend from the hub. The row includes 12 (N) of the fan blades, a solidity value (R) that is from 1.0 to 1.2, and a ratio of N/R that is from 10.0 to 12.0.

CROSS REFERENCE TO RELATED APPLICATIONS

The present disclosure is a continuation of U.S. application Ser. No.13/176,365, filed Jul. 5, 2011.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

This invention was made with government support under contract numberNAS3-01138 awarded by NASA. The government has certain rights in theinvention.

BACKGROUND

This disclosure relates to gas turbine engines and, more particularly,to an engine having a geared turbo fan architecture that is designed toefficiently operate with a high bypass ratio and a low pressure ratio.

The propulsive efficiency of a gas turbine engine depends on manydifferent factors, such as the design of the engine and the resultingperformance debits on the fan that propels the engine. As an example,the fan rotates at a high rate of speed such that air passes over theblades at transonic or supersonic speeds. The fast-moving air createsflow discontinuities or shocks that result in irreversible propulsivelosses. Additionally, physical interaction between the fan and the aircauses downstream turbulence and further losses. Although some basicprinciples behind such losses are understood, identifying and changingappropriate design factors to reduce such losses for a given enginearchitecture has proven to be a complex and elusive task.

SUMMARY

An exemplary gas turbine engine includes a spool, a turbine coupled todrive the spool, and a propulsor that is coupled to be driven by theturbine through the spool. A gear assembly is coupled between thepropulsor and the spool such that rotation of the turbine drives thepropulsor at a different speed than the spool. The propulsor includes ahub and a row of propulsor blades that extend from the hub. The rowincludes no more than 20 of the propulsor blades.

In another aspect, a gas turbine engine includes a core flow passage anda bypass flow passage. A propulsor is arranged at an inlet of the bypassflow passage and core flow passage. The propulsor includes a hub and arow of propulsor blades that extend from the hub. The row includes nomore than 20 of the propulsor blades and the bypass flow passage has adesign pressure ratio of approximately 1.3-1.55 with regard to an inletpressure and an outlet pressure of the bypass flow passage.

An exemplary propulsor for use in a gas turbine engine includes a rotorhaving a row of propulsor blades that extends radially outwardly from ahub. Each of the propulsor blades extends radially between a root and atip and in a chord direction between a leading edge and a trailing edgeto define a chord dimension at the tip of each propulsor blade. The rowof propulsor blades defines a circumferential pitch with regard to thetips. The row of propulsor blades has a solidity value defined as thechord dimension divided by the circumferential pitch. The row alsoincludes a number of the propulsor blades that is no greater than 20such that a ratio of the number of propulsor blades to the solidityvalue is from 9 to 20.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of the disclosed examples willbecome apparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

FIG. 1 is a schematic cross-section of a gas turbine engine.

FIG. 2 is a perspective view of a fan section of the engine of FIG. 1.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 may be a two-spool turbofan that generallyincorporates a fan section 22, a compressor section 24, a combustorsection 26 and a turbine section 28. Alternative engine architecturesmay include a single-spool design, a three-spool design, or an openrotor design, among other systems or features.

The fan section 22 drives air along a bypass flow passage B while thecompressor section 24 drives air along a core flow passage C forcompression and communication into the combustor section 26. Althoughdepicted as a turbofan gas turbine engine, it is to be understood thatthe concepts described herein are not limited to use with turbofans andthe teachings may be applied to other types of gas turbine engines.

The engine 20 includes a low speed spool 30 and high speed spool 32mounted for rotation about an engine central longitudinal axis Arelative to an engine static structure 36 via several bearing systems38. The low speed spool 30 generally includes an inner shaft 40 that iscoupled with a propulsor 42, a low pressure compressor 44 and a lowpressure turbine 46. The low pressure turbine 46 drives the propulsor 42through the inner shaft 40 and a gear assembly 48, which allows the lowspeed spool 30 to drive the propulsor 42 at a different (e.g. lower)angular speed.

The high speed spool 32 includes an outer shaft 50 that is coupled witha high pressure compressor 52 and a high pressure turbine 54. Acombustor 56 is arranged between the high pressure compressor 52 and thehigh pressure turbine 54. The inner shaft 40 and the outer shaft 50 areconcentric and rotate about the engine central longitudinal axis A,which is collinear with their longitudinal axes.

A core airflow in core flow passage C is compressed by the low pressurecompressor 44 then the high pressure compressor 52, mixed with the fuelin the combustor 56, and then expanded over the high pressure turbine 54and low pressure turbine 46. The turbines 54, 46 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

As shown, the propulsor 42 is arranged at an inlet 60 of the bypass flowpassage B and core flow passage C. Air flow through the bypass flowpassage B exits the engine 20 through an outlet 62 or nozzle. For agiven design of the propulsor 42, the inlet 60 and the outlet 62 of theengine 20 define a design pressure ratio with regard to an inletpressure at the inlet 60 and an outlet pressure at the outlet 62 of thebypass flow passage B. As an example, the design pressure ratio may bedetermined based upon the stagnation inlet pressure and the stagnationoutlet pressure at a design rotational speed of the engine 20. In thatregard, the engine 20 may optionally include a variable area nozzle 64within the bypass flow passage B. The variable area nozzle 64 isoperative to change a cross-sectional area 66 of the outlet 62 tothereby control the pressure ratio via changing pressure within thebypass flow passage B. The design pressure ratio may be defined with thevariable area nozzle 64 fully open or fully closed.

Referring to FIG. 2, the propulsor 42, which in this example is a fan,includes a rotor 70 having a row 72 of propulsor blades 74 that extend acircumferentially around a hub 76. Each of the propulsor blades 74extends radially outwardly from the hub 76 between a root 78 and a tip80 and in a chord direction (axially and circumferentially) between aleading edge 82 and a trailing edge 84. A chord dimension (CD) is alength between the leading edge 82 and the trailing edge 84 at the tipof each propulsor blade 74. The row 72 of propulsor blades 74 alsodefines a circumferential pitch (CP) that is equivalent to the arcdistance between the tips 80 of neighboring propulsor blades 74.

As will be described, the example propulsor 42 includes a number (N) ofthe propulsor blades 74 and a geometry that, in combination with thearchitecture of the engine 20, provides enhanced propulsive efficiencyby reducing performance debits of the propulsor 42.

In the illustrated example, the number N of propulsor blades in the row72 is no more than 20. In one example, the propulsor 42 includes 18 ofthe propulsor blades 74 uniformly circumferentially arranged about thehub 76. In other embodiments, the number N may be any number of bladesfrom 12-20.

The propulsor blades 74 define a solidity value with regard to the chorddimension CD and the circumferential pitch CP. The solidity value isdefined as a ratio (R) of CD/CP (i.e., CD divided by CP). Inembodiments, the solidity value of the propulsor 42 is between 1.0 and1.3. In further embodiments, the solidity value is from 1.1 to 1.2.

Additionally, in combination with the given example solidity values, theengine 20 may be designed with a particular design pressure ratio. Inembodiments, the design pressure ratio may be between 1.3 and 1.55. In afurther embodiment, the design pressure ratio may be between 1.3 and1.4.

The engine 20 may also be designed with a particular bypass ratio withregard to the amount of air that passes through the bypass flow passageB and the amount of air that passes through the core flow passage C. Asan example, the design bypass ratio of the engine 20 may nominally be12, or alternatively in a range of approximately 8.5 to 13.5.

The propulsor 42 also defines a ratio of N/R. In embodiments, the ratioN/R is from 9 to 20. In further embodiments, the ratio N/R is from 14 to16. The table below shows additional examples of solidity and the ratioN/R for different numbers of propulsor blades 74.

TABLE Number of Blades, Solidity and Ratio N/R Number of Blades (N)Solidity Ratio N/R 20 1.3 15.4 18 1.3 13.8 16 1.3 12.3 14 1.3 10.8 121.3 9.2 20 1.2 16.7 18 1.2 15.0 16 1.2 13.3 14 1.2 11.7 12 1.2 10.0 201.1 18.2 18 1.1 16.4 16 1.1 14.5 14 1.1 12.7 12 1.1 10.9 20 1.0 20.0 181.0 18.0 16 1.0 16.0 14 1.0 14.0 12 1.0 12.0

The disclosed ratios of N/R enhance the propulsive efficiency of thedisclosed engine 20. For instance, the disclosed ratios of N/R aredesigned for the geared turbo fan architecture of the engine 20 thatutilizes the gear assembly 48. That is, the gear assembly 48 allows thepropulsor 42 to rotate at a different, lower speed than the low speedspool 30. In combination with the variable area nozzle 64, the propulsor42 can be designed with a large diameter and rotate at a relatively slowspeed with regard to the low speed spool 30. A relatively low speed,relatively large diameter, and the geometry that permits the disclosedratios of N/R contribute to the reduction of performance debits, such asby lowering the speed of the air or fluid that passes over the propulsorblades 74.

Although a combination of features is shown in the illustrated examples,not all of them need to be combined to realize the benefits of variousembodiments of this disclosure. In other words, a system designedaccording to an embodiment of this disclosure will not necessarilyinclude all of the features shown in any one of the Figures or all ofthe portions schematically shown in the Figures. Moreover, selectedfeatures of one example embodiment may be combined with selectedfeatures of other example embodiments.

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthe essence of this disclosure. The scope of legal protection given tothis disclosure can only be determined by studying the following claims.

What is claimed is:
 1. A gas turbine engine comprising: a gear assembly;a bypass flow passage and a core flow passage, the bypass flow passageincluding an inlet; a fan arranged within the bypass flow passage; afirst shaft and a second shaft; a first turbine coupled with the firstshaft, the first shaft coupled through the gear assembly with the fan;wherein the fan includes a hub and a row of fan blades that extend fromthe hub, and the row includes a number (N) of the fan blades, the number(N) being 12, a solidity value (R) at tips of the fan blades that isfrom 1.0 to 1.2, and a ratio of N/R that is from 10.0 to 12.0.
 2. Thegas turbine engine as recited in claim 1, further comprising a secondturbine coupled with the second shaft, wherein the second turbine is a2-stage turbine.
 3. The gas turbine engine as recited in claim 1,wherein the bypass flow passage includes an outlet, the inlet and theoutlet defining a design pressure ratio with regard to an inlet pressureat the inlet and an outlet pressure at the outlet at a design rotationalspeed of the engine, the design pressure ratio being approximately 1.3to 1.55.
 4. The gas turbine engine as recited in claim 3, wherein thedesign pressure ratio is between 1.3 and 1.4.
 5. The gas turbine engineas recited in claim 4, further comprising a second turbine coupled withthe second shaft, wherein the second turbine is a 2-stage turbine. 6.The gas turbine engine as recited in claim 5, wherein the ratio of N/Ris from 10.9 to 12.0.
 7. The gas turbine engine as recited in claim 1,wherein the ratio of N/R is from 10.9 to 12.0.
 8. The gas turbine engineas recited in claim 7, further comprising a second turbine coupled withthe second shaft, wherein the second turbine is a 2-stage turbine. 9.The gas turbine engine as recited in claim 8, wherein the bypass flowpassage includes an outlet, the inlet and the outlet defining a designpressure ratio with regard to an inlet pressure at the inlet and anoutlet pressure at the outlet at a design rotational speed of theengine, the design pressure ratio being approximately 1.3 to 1.55. 10.The gas turbine engine as recited in claim 9, wherein the designpressure ratio is between 1.3 and 1.4.
 11. A gas turbine enginecomprising: a gear assembly; a bypass flow passage and a core flowpassage, the bypass flow passage including an inlet an an outlet, theinlet and the outlet defining a design pressure ratio with regard to aninlet pressure at the inlet and an outlet pressure at the outlet at adesign rotational speed of the engine, the design pressure ratio beingapproximately 1.3 to 1.4; a fan arranged within the bypass flow passage;a first shaft and a second shaft; a first turbine coupled with the firstshaft, the first shaft coupled through the gear assembly with the fan;wherein the fan includes a hub and a row of fan blades that extend fromthe hub, and the row includes a number (N) of the fan blades, the number(N) being 16, a solidity value (R) at tips of the fan blades that isfrom 1.0 to 1.2, and a ratio of N/R that is from 13.3 to 16.0.
 12. Thegas turbine engine as recited in claim 11, further comprising a secondturbine coupled with the second shaft, wherein the second turbine is a2-stage turbine.
 13. The gas turbine engine as recited in claim 12,further comprising a first compressor coupled with the first shaft,wherein the first compressor is a 3-stage compressor.
 14. The gasturbine engine as recited in claim 11, further comprising a secondturbine coupled with the second shaft, wherein the second turbine is a2-stage turbine.
 15. The gas turbine engine as recited in claim 14,wherein the ratio of N/R is from 14.5 to 16.0.
 16. The gas turbineengine as recited in claim 11, wherein the ratio of N/R is from 14.5 to16.0.
 17. The gas turbine engine as recited in claim 16, wherein thesolidity value (R) at the tips of the fan blades is from 1.0 to 1.1. 18.The gas turbine engine as recited in claim 17, further comprising asecond turbine coupled with the second shaft, wherein the second turbineis a 2-stage turbine.
 19. A gas turbine engine comprising: a gearassembly; a bypass flow passage and a core flow passage, the bypass flowpassage including an inlet; a fan arranged within the bypass flowpassage; a first shaft and a second shaft; a first turbine coupled withthe first shaft, the first shaft coupled through the gear assembly withthe fan; wherein the fan includes a hub and a row of fan blades thatextend from the hub, and the row includes a number (N) of the fanblades, the number (N) being 14, a solidity value (R) at tips of the fanblades that is from 1.0 to 1.2, and a ratio of N/R that is from 11.7 to14.0.
 20. The gas turbine engine as recited in claim 19, furthercomprising a second turbine coupled with the second shaft, wherein thesecond turbine is a 2-stage turbine.
 21. The gas turbine engine asrecited in claim 19, wherein the bypass flow passage includes an outlet,the inlet and the outlet defining a design pressure ratio with regard toan inlet pressure at the inlet and an outlet pressure at the outlet at adesign rotational speed of the engine, the design pressure ratio beingapproximately 1.3 to 1.55.
 22. The gas turbine engine as recited inclaim 21, further comprising a second turbine coupled with the secondshaft, wherein the second turbine is a 2-stage turbine.
 23. The gasturbine engine as recited in claim 22, wherein the ratio of N/R is from12.7 to 14.0.
 24. The gas turbine engine as recited in claim 19, whereinthe ratio of N/R is from 12.7 to 14.0.
 25. The gas turbine engine asrecited in claim 24, further comprising a second turbine coupled withthe second shaft, wherein the second turbine is a 2-stage turbine. 26.The gas turbine engine as recited in claim 25, wherein the bypass flowpassage includes an outlet, the inlet and the outlet defining a designpressure ratio with regard to an inlet pressure at the inlet and anoutlet pressure at the outlet at a design rotational speed of theengine, the design pressure ratio being between 1.3 and 1.4.
 27. A gasturbine engine comprising: a bypass flow passage and a core flowpassage, the bypass flow passage including an inlet and an outlet whichdefine a design pressure ratio with regard to an inlet pressure at theinlet and an outlet pressure at the outlet at a design rotational speedof the engine, the design pressure ratio being between 1.3 and 1.4; afan arranged within the bypass flow passage; a first shaft and a secondshaft; a first turbine coupled with the first, the first shaft coupledwith the fan; and a second turbine coupled with a second shaft; whereinthe fan includes a hub and a row of fan blades that extend from the hub,and the row includes a number (N) of the fan blades that is from 12 to16, a solidity value (R) at tips of the fan blades that is from 1.0 to1.2, and a ratio of N/R that is from 10.0 to 16.0.
 28. The gas turbineengine as recited in claim 27, further comprising a gear assembly,wherein the first turbine is coupled with the fan through the firstshaft and gear assembly, and further comprising a variable area nozzle,wherein the design pressure ratio is achieved in operation with thevisible area nozzle fully open.